Gas turbine engine with low stage count low pressure turbine

ABSTRACT

A gas turbine engine includes a spool along an engine centerline axis which drives a gear train, said spool includes a low stage count low pressure turbine.

CROSS REFERENCE TO RELATED APPLICATIONS

The present disclosure is a continuation in part of U.S. patentapplication Ser. No. 12/131876, filed Jun. 2, 2008.

BACKGROUND

The present invention relates to a gas turbine engine and moreparticularly to an engine mounting configuration for the mounting of aturbofan gas turbine engine to an aircraft pylon.

A gas turbine engine may be mounted at various points on an aircraftsuch as a pylon integrated with an aircraft structure. An enginemounting configuration ensures the transmission of loads between theengine and the aircraft structure. The loads typically include theweight of the engine, thrust, aerodynamic side loads, and rotary torqueabout the engine axis. The engine mount configuration must also absorbthe deformations to which the engine is subjected during differentflight phases and the dimensional variations due to thermal expansionand retraction.

One conventional engine mounting configuration includes a pylon having aforward mount and an aft mount with relatively long thrust links whichextend forward from the aft mount to the engine intermediate casestructure. Although effective, one disadvantage of this conventionaltype mounting arrangement is the relatively large “punch loads” into theengine cases from the thrust links which react the thrust from theengine and couple the thrust to the pylon. These loads tend to distortthe intermediate case and the low pressure compressor (LPC) cases. Thedistortion may cause the clearances between the static cases androtating blade tips to increase which may negatively affect engineperformance and increase fuel burn.

SUMMARY

A gas turbine engine according to an exemplary aspect of the presentdisclosure includes a gear train defined along an engine centerlineaxis, and a spool along said engine centerline axis which drives thegear train, the spool includes a low stage count low pressure turbine.

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the low stage count may include three to six (3-6)stages. Additionally or alternatively, the low stage count may includethree (3) stages. Additionally or alternatively, the low stage count mayinclude five (5) or six (6) stages.

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the spool may be a low spool.

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the engine may include a core nacelle defined aboutthe engine centerline axis, a fan nacelle mounted at least partiallyaround the core nacelle to define a fan bypass flow path for a fanbypass airflow, and a fan variable area nozzle axially movable relativethe fan nacelle to vary a fan nozzle exit area and adjust a pressureratio of the fan bypass airflow during engine operation.

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the engine may include a controller operable tocontrol the fan variable area nozzle to vary a fan nozzle exit area andadjust the pressure ratio of the fan bypass airflow.

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the controller may be operable to reduce the fannozzle exit area at a cruise flight condition. Additionally oralternatively, the controller may be operable to control the fan nozzleexit area to reduce a fan instability.

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the fan variable area nozzle may define a trailingedge of the fan nacelle.

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the bypass flow may define a bypass ratio greaterthan about six (6). Additionally or alternatively, the bypass flow maydefine a bypass ratio greater than about ten (10). Additionally oralternatively, the bypass flow may define a bypass ratio greater thansix (6). Additionally or alternatively, the bypass flow may define abypass ratio greater than ten (10).

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the gear train may define a gear reduction ratio ofgreater than or equal to about 2.3. Additionally or alternatively, thegear train may define a gear reduction ratio of greater than or equal toabout 2.5. Additionally or alternatively, the gear train may define agear reduction ratio of greater than or equal to 2.5.

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the low pressure turbine may define a pressure ratiothat is greater than about five (5).

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the low pressure turbine may define a pressure ratiothat is greater than five (5).

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the gear train may drive a fan.

A gas turbine engine according to another exemplary aspect of thepresent disclosure includes a core nacelle defined about an enginecenterline axis, a fan nacelle mounted at least partially around thecore nacelle to define a fan bypass flow path for a fan bypass airflow,a gear train within the core nacelle, a spool along the enginecenterline axis within the core nacelle to drive the gear train, thespool includes a low stage count low pressure turbine, and a fanvariable area nozzle axially movable relative to the fan nacelle to varya fan nozzle exit area and adjust a pressure ratio of the fan bypassairflow during engine operation.

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the low pressure turbine may define a pressure ratiothat is greater than five (5).

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the low pressure turbine may define a pressure ratiothat is greater than five (5), the bypass flow defines a bypass ratiogreater than ten (10), and the gear train defines a gear reduction ratioof greater than or equal to 2.5.

BRIEF DESCRIPTION OF THE DRAWINGS

The various features and advantages of this invention will becomeapparent to those skilled in the art from the following detaileddescription of the currently disclosed embodiment. The drawings thataccompany the detailed description can be briefly described as follows:

FIG. 1A is a general schematic sectional view through a gas turbineengine along the engine longitudinal axis;

FIG. 1B is a general sectional view through a gas turbine engine alongthe engine longitudinal axis illustrating an engine static structurecase arrangement on the lower half thereof;

FIG. 1C is a side view of an mount system illustrating a rear mountattached through an engine thrust case to a mid-turbine frame between afirst and second bearing supported thereby;

FIG. 1D is a forward perspective view of an mount system illustrating arear mount attached through an engine thrust case to a mid-turbine framebetween a first and second bearing supported thereby;

FIG. 2A is a top view of an engine mount system;

FIG. 2B is a side view of an engine mount system within a nacellesystem;

FIG. 2C is a forward perspective view of an engine mount system within anacelle system;

FIG. 3 is a side view of an engine mount system within another frontmount;

FIG. 4A is an aft perspective view of an aft mount;

FIG. 4B is an aft view of an aft mount of FIG. 4A;

FIG. 4C is a front view of the aft mount of FIG. 4A;

FIG. 4D is a side view of the aft mount of FIG. 4A;

FIG. 4E is a top view of the aft mount of FIG. 4A;

FIG. 5A is a side view of the aft mount of FIG. 4A in a first slideposition; and

FIG. 5B is a side view of the aft mount of FIG. 4A in a second slideposition.

DETAILED DESCRIPTION OF THE DISCLOSED EMBODIMENT

FIG. 1A illustrates a general partial fragmentary schematic view of agas turbofan engine 10 suspended from an engine pylon 12 within anengine nacelle assembly N as is typical of an aircraft designed forsubsonic operation.

The turbofan engine 10 includes a core engine within a core nacelle Cthat houses a low spool 14 and high spool 24. The low spool 14 includesa low pressure compressor 16 and low pressure turbine 18. The low spool14 drives a fan section 20 connected to the low spool 14 either directlyor through a gear train 25.

The high spool 24 includes a high pressure compressor 26 and highpressure turbine 28. A combustor 30 is arranged between the highpressure compressor 26 and high pressure turbine 28. The low and highspools 14, 24 rotate about an engine axis of rotation A.

The engine 10 in one non-limiting embodiment is a high-bypass gearedarchitecture aircraft engine. In one disclosed, non-limiting embodiment,the engine 10 bypass ratio is greater than about six (6), with anexample embodiment being greater than about ten (10), the gear train 25is an epicyclic gear train such as a planetary gear system or other gearsystem with a gear reduction ratio of greater than about 2.3 and the lowpressure turbine 18 has a pressure ratio that is greater than about 5.In one disclosed embodiment, the engine 10 bypass ratio is greater thanten (10:1), the turbofan diameter is significantly larger than that ofthe low pressure compressor 16, and the low pressure turbine 18 has apressure ratio that is greater than 5:1. The gear train 25 may be anepicycle gear train such as a planetary gear system or other gear systemwith a gear reduction ratio of greater than about 2.5:1. It should beunderstood, however, that the above parameters are only exemplary of oneembodiment of a geared architecture engine and that the presentinvention is applicable to other gas turbine engines including directdrive turbofans.

Airflow enters the fan nacelle F which at least partially surrounds thecore nacelle C. The fan section 20 communicates airflow into the corenacelle C to the low pressure compressor 16. Core airflow compressed bythe low pressure compressor 16 and the high pressure compressor 26 ismixed with the fuel in the combustor 30 where is ignited, and burned.The resultant high pressure combustor products are expanded through thehigh pressure turbine 28 and low pressure turbine 18. The turbines 28,18 are rotationally coupled to the compressors 26, 16 respectively todrive the compressors 26, 16 in response to the expansion of thecombustor product. The low pressure turbine 18 also drives the fansection 20 through gear train 25. A core engine exhaust E exits the corenacelle C through a core nozzle 43 defined between the core nacelle Cand a tail cone 33.

With reference to FIG. 1B, the low pressure turbine 18 includes a lownumber of stages, which, in the illustrated non-limiting embodiment,includes three turbine stages, 18A, 18B, 18C. The gear train 22operationally effectuates the significantly reduced number of stageswithin the low pressure turbine 18. The three turbine stages, 18A, 18B,18C facilitate a lightweight and operationally efficient enginearchitecture. It should be appreciated that a low number of stagescontemplates, for example, three to six (3-6) stages. Low pressureturbine 18 pressure ratio is pressure measured prior to inlet of lowpressure turbine 18 as related to the pressure at the outlet of the lowpressure turbine 18 prior to exhaust nozzle.

Thrust is a function of density, velocity, and area. One or more ofthese parameters can be manipulated to vary the amount and direction ofthrust provided by the bypass flow B. The Variable Area Fan Nozzle(“VAFN”) 42 operates to effectively vary the area of the fan nozzle exitarea 44 to selectively adjust the pressure ratio of the bypass flow B inresponse to a controller C. Low pressure ratio turbofans are desirablefor their high propulsive efficiency. However, low pressure ratio fansmay be inherently susceptible to fan stability/flutter problems at lowpower and low flight speeds. The VAFN 42 allows the engine to change toa more favorable fan operating line at low power, avoiding theinstability region, and still provide the relatively smaller nozzle areanecessary to obtain a high-efficiency fan operating line at cruise.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 20 of the engine 10 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, withthe engine at its best fuel consumption—also known as “bucket cruiseThrust Specific Fuel Consumption (‘TSFCT’)”—is the industry standardparameter of 1bm of fuel being burned divided by 1bf of thrust theengine produces at that minimum point. “Low fan pressure ratio” is thepressure ratio across the fan blade alone, without the Fan Exit GuideVane (“FEGV”) system 36. The low fan pressure ratio as disclosed hereinaccording to one non-limiting embodiment is less than about 1.45. “Lowcorrected fan tip speed” is the actual fan tip speed in ft/sec dividedby an industry standard temperature correction of [(Tambient degR)/518.7)̂0.5]. The “Low corrected fan tip speed” as disclosed hereinaccording to one non-limiting embodiment is less than about 1150ft/second.

As the fan blades within the fan section 20 are efficiently designed ata particular fixed stagger angle for an efficient cruise condition, theVAFN 42 is operated to effectively vary the fan nozzle exit area 44 toadjust fan bypass air flow such that the angle of attack or incidence onthe fan blades is maintained close to the design incidence for efficientengine operation at other flight conditions, such as landing and takeoffto thus provide optimized engine operation over a range of flightconditions with respect to performance and other operational parameterssuch as noise levels.

The engine static structure 44 generally has sub-structures including acase structure often referred to as the engine backbone. The enginestatic structure 44 generally includes a fan case 46, an intermediatecase (IMC) 48, a high pressure compressor case 50, a combustor case 52A,a high pressure turbine case 52B, a thrust case 52C, a low pressureturbine case 54, and a turbine exhaust case 56 (FIG. 1B). Alternatively,the combustor case 52A, the high pressure turbine case 52B and thethrust case 52C may be combined into a single case. It should beunderstood that this is an exemplary configuration and any number ofcases may be utilized.

The fan section 20 includes a fan rotor 32 with a plurality ofcircumferentially spaced radially outwardly extending fan blades 34. Thefan blades 34 are surrounded by the fan case 46. The core engine casestructure is secured to the fan case 46 at the IMC 48 which includes amultiple of circumferentially spaced radially extending struts 40 whichradially span the core engine case structure and the fan case 20.

The engine static structure 44 further supports a bearing system uponwhich the turbines 28, 18, compressors 26, 16 and fan rotor 32 rotate. A#1 fan dual bearing 60 which rotationally supports the fan rotor 32 isaxially located generally within the fan case 46. The #1 fan dualbearing 60 is preloaded to react fan thrust forward and aft (in case ofsurge). A #2 LPC bearing 62 which rotationally supports the low spool 14is axially located generally within the intermediate case (IMC) 48. The#2 LPC bearing 62 reacts thrust. A #3 fan dual bearing 64 whichrotationally supports the high spool 24 and also reacts thrust. The #3fan bearing 64 is also axially located generally within the IMC 48 justforward of the high pressure compressor case 50. A #4 bearing 66 whichrotationally supports a rear segment of the low spool 14 reacts onlyradial loads. The #4 bearing 66 is axially located generally within thethrust case 52C in an aft section thereof. A #5 bearing 68 rotationallysupports the rear segment of the low spool 14 and reacts only radialloads. The #5 bearing 68 is axially located generally within the thrustcase 52C just aft of the #4 bearing 66. It should be understood thatthis is an exemplary configuration and any number of bearings may beutilized.

The #4 bearing 66 and the #5 bearing 68 are supported within amid-turbine frame (MTF) 70 to straddle radially extending structuralstruts 72 which are preloaded in tension (FIGS. 1C-1D). The MTF 70provides aft structural support within the thrust case 52C for the #4bearing 66 and the #5 bearing 68 which rotatably support the spools 14,24.

A dual rotor engine such as that disclosed in the illustrated embodimenttypically includes a forward frame and a rear frame that support themain rotor bearings. The intermediate case (IMC) 48 also includes theradially extending struts 40 which are generally radially aligned withthe #2 LPC bearing 62 (FIG. 1B). It should be understood that variousengines with various case and frame structures will benefit from thepresent invention.

The turbofan gas turbine engine 10 is mounted to aircraft structure suchas an aircraft wing through a mount system 80 attachable by the pylon12. The mount system 80 includes a forward mount 82 and an aft mount 84(FIG. 2A). The forward mount 82 is secured to the IMC 48 and the aftmount 84 is secured to the MTF 70 at the thrust case 52C. The forwardmount 82 and the aft mount 84 are arranged in a plane containing theaxis A of the turbofan gas turbine 10. This eliminates the thrust linksfrom the intermediate case, which frees up valuable space beneath thecore nacelle and minimizes IMC 48 distortion.

Referring to FIGS. 2A-2C, the mount system 80 reacts the engine thrustat the aft end of the engine 10. The term “reacts” as utilized in thisdisclosure is defined as absorbing a load and dissipating the load toanother location of the gas turbine engine 10.

The forward mount 82 supports vertical loads and side loads. The forwardmount 82 in one non-limiting embodiment includes a shackle arrangementwhich mounts to the IMC 48 at two points 86A, 86B. The forward mount 82is generally a plate-like member which is oriented transverse to theplane which contains engine axis A. Fasteners are oriented through theforward mount 82 to engage the intermediate case (IMC) 48 generallyparallel to the engine axis A. In this illustrated non-limitingembodiment, the forward mount 82 is secured to the IMC 40. In anothernon-limiting embodiment, the forward mount 82 is secured to a portion ofthe core engine, such as the high-pressure compressor case 50 of the gasturbine engine 10 (see FIG. 3). One of ordinary skill in the art havingthe benefit of this disclosure would be able to select an appropriatemounting location for the forward mount 82.

Referring to FIG. 4A, the aft mount 84 generally includes a first A-arm88A, a second A-arm 88B, a rear mount platform 90, a wiffle treeassembly 92 and a drag link 94. The rear mount platform 90 is attacheddirectly to aircraft structure such as the pylon 12. The first A-arm 88Aand the second A-arm 88B mount between the thrust case 52C at casebosses 96 which interact with the MTF 70 (FIGS. 4B-4C), the rear mountplatform 90 and the wiffle tree assembly 92. It should be understoodthat the first A-arm 88A and the second A-arm 88B may alternativelymount to other areas of the engine 10 such as the high pressure turbinecase or other cases. It should also be understood that other framearrangements may alternatively be used with any engine case arrangement.

Referring to FIG. 4D, the first A-arm 88A and the second A-arm 88B arerigid generally triangular arrangements, each having a first link arm 89a, a second link arm 89 b and a third link arm 89 c. The first link arm89 a is between the case boss 96 and the rear mount platform 90. Thesecond link arm 89 b is between the case bosses 96 and the wiffle treeassembly 92. The third link arm 89 c is between the wiffle tree assembly92 rear mount platform 90. The first A-arm 88A and the second A-arm 88Bprimarily support the vertical weight load of the engine 10 and transmitthrust loads from the engine to the rear mount platform 90.

The first A-arm 88A and the second A-arm 88B of the aft mount 84 forcethe resultant thrust vector at the engine casing to be reacted along theengine axis A which minimizes tip clearance losses due to engine loadingat the aft mount 84. This minimizes blade tip clearance requirements andthereby improves engine performance.

The wiffle tree assembly 92 includes a wiffle link 98 which supports acentral ball joint 100, a first sliding ball joint 102A and a secondsliding ball joint 102B (FIG. 4E). It should be understood that variousbushings, vibration isolators and such like may additionally be utilizedherewith. The central ball joint 100 is attached directly to aircraftstructure such as the pylon 12. The first sliding ball joint 102A isattached to the first A-arm 88A and the second sliding ball joint 102Bis mounted to the first A-arm 88A. The first and second sliding balljoint 102A, 102B permit sliding movement of the first and second A-arm88A, 88B (illustrated by arrow S in FIGS. 5A and 5B) to assure that onlya vertical load is reacted by the wiffle tree assembly 92. That is, thewiffle tree assembly 92 allows all engine thrust loads to be equalizedtransmitted to the engine pylon 12 through the rear mount platform 90 bythe sliding movement and equalize the thrust load that results from thedual thrust link configuration. The wiffle link 98 operates as anequalizing link for vertical loads due to the first sliding ball joint102A and the second sliding ball joint 102B. As the wiffle link 98rotates about the central ball joint 100 thrust forces are equalized inthe axial direction. The wiffle tree assembly 92 experiences loadingonly due to vertical loads, and is thus less susceptible to failure thanconventional thrust-loaded designs.

The drag link 94 includes a ball joint 104A mounted to the thrust case52C and ball joint 104B mounted to the rear mount platform 90 (FIGS.4B-4C). The drag link 94 operates to react torque.

The aft mount 84 transmits engine loads directly to the thrust case 52Cand the MTF 70. Thrust, vertical, side, and torque loads are transmitteddirectly from the MTF 70 which reduces the number of structural membersas compared to current in-practice designs.

The mount system 80 is compact, and occupies space within the corenacelle volume as compared to turbine exhaust case-mountedconfigurations, which occupy space outside of the core nacelle which mayrequire additional or relatively larger aerodynamic fairings andincrease aerodynamic drag and fuel consumption. The mount system 80eliminates the heretofore required thrust links from the IMC, whichfrees up valuable space adjacent the IMC 48 and the high pressurecompressor case 50 within the core nacelle C.

It should be understood that relative positional terms such as“forward,” “aft,” “upper,” “lower,” “above,” “below,” and the like arewith reference to the normal operational attitude of the vehicle andshould not be considered otherwise limiting.

The foregoing description is exemplary rather than defined by thelimitations within. Many modifications and variations of the presentinvention are possible in light of the above teachings. The disclosedembodiments of this invention have been disclosed, however, one ofordinary skill in the art would recognize that certain modificationswould come within the scope of this invention. It is, therefore, to beunderstood that within the scope of the appended claims, the inventionmay be practiced otherwise than as specifically described. For thatreason the following claims should be studied to determine the truescope and content of this invention.

1. A gas turbine engine comprising: a gear train defined along an enginecenterline axis; and a spool along said engine centerline axis whichdrives said gear train, said spool includes a low stage count lowpressure turbine.
 2. The engine as recited in claim 1, wherein said lowstage count includes three to six (3-6) stages.
 3. The engine as recitedin claim 1, wherein said low stage count includes three (3) stages. 4.The engine as recited in claim 1, wherein said low stage count includesfive (5) stages.
 5. The engine as recited in claim 1, wherein said lowstage count includes six (6) stages.
 6. The engine as recited in claim1, wherein said spool is a low spool.
 7. The engine as recited in claim1, further comprising: a core nacelle defined about said enginecenterline axis; a fan nacelle mounted at least partially around saidcore nacelle to define a fan bypass airflow flow path for a fan bypassairflow; and a fan variable area nozzle axially movable relative saidfan nacelle to vary a fan nozzle exit area and adjust a pressure ratioof the fan bypass airflow during engine operation.
 8. The engine asrecited in claim 7, further comprising: a controller operable to controlsaid fan variable area nozzle to vary the fan nozzle exit area andadjust the pressure ratio of the fan bypass airflow.
 9. The engine asrecited in claim 8, wherein said controller is operable to reduce saidfan nozzle exit area at a cruise flight condition.
 10. The engine asrecited in claim 8, wherein said controller is operable to control saidfan nozzle exit area to reduce a fan instability.
 11. The engine asrecited in claim 8, wherein said fan variable area nozzle defines atrailing edge of said fan nacelle.
 12. The engine as recited in claim 8,wherein said fan bypass airflow defines a bypass ratio greater thanabout six (6).
 13. The engine as recited in claim 8, wherein said fanbypass airflow defines a bypass ratio greater than about ten (10). 14.The engine as recited in claim 8, wherein said fan bypass airflowdefines a bypass ratio greater than six (6).
 15. The engine as recitedin claim 8, wherein said fan bypass airflow defines a bypass ratiogreater than ten (10).
 16. The engine as recited in claim 1, whereinsaid gear train defines a gear reduction ratio of greater than or equalto about 2.3.
 17. The engine as recited in claim 1, wherein said geartrain defines a gear reduction ratio of greater than or equal to about2.5.
 18. The engine as recited in claim 1, wherein said gear traindefines a gear reduction ratio of greater than or equal to 2.5.
 19. Theengine as recited in claim 1, wherein said low pressure turbine definesa low pressure turbine pressure ratio that is greater than about five(5).
 20. The engine as recited in claim 1, wherein said low pressureturbine defines a low pressure turbine pressure ratio that is greaterthan five (5).
 21. The engine as recited in claim 1, wherein said spooldrives a fan through said gear train.
 22. A gas turbine enginecomprising: a core nacelle defined about an engine centerline axis; afan nacelle mounted at least partially around said core nacelle todefine a fan bypass flow path for a fan bypass airflow; a gear trainwithin said core nacelle; a spool along said engine centerline axiswithin said core nacelle to drive said gear train, said spool includes alow stage count low pressure turbine; and a fan variable area nozzleaxially movable relative to said fan nacelle to vary a fan nozzle exitarea and adjust a pressure ratio of the fan bypass airflow during engineoperation.
 23. The engine as recited in claim 22, wherein said lowpressure turbine defines a pressure ratio that is greater than five (5).24. The engine as recited in claim 22, wherein said low pressure turbinedefines a pressure ratio that is greater than five (5), said spooldrives a fan within said fan nacelle through said gear train, said fanbypass airflow defines a bypass ratio greater than ten (10), and saidgear train defines a gear reduction ratio of greater than or equal to2.5.